Dual function cascade integrated variable area fan nozzle and thrust reverser

ABSTRACT

A gas turbine engine system according to an exemplary aspect of the present disclosure may include a core engine defined about an axis, a fan driven by the core engine about the axis to generate bypass flow, and at least one integrated mechanism in communication with the bypass flow. The at least one integrated mechanism includes a variable area fan nozzle (VAFN) and thrust reverser, and a plurality of positions to control bypass flow.”

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.15/674,576 filed Aug. 11, 2017, which is a continuation of U.S.application Ser. No. 13/332,529 filed Dec. 21, 2011, which is acontinuation-in-part of U.S. application Ser. No. 12/440,746 filed Mar.11, 2009, which is a National Phase application of PCT/US2006/039990filed Oct. 12, 2006.

BACKGROUND

This invention relates to gas turbine engines and, more particularly, toa gas turbine engine having a variable fan nozzle integrated with athrust reverser of the gas turbine engine.

Gas turbine engines are widely known and used for power generation andvehicle (e.g., aircraft) propulsion. A typical gas turbine engineincludes a compression section, a combustion section, and a turbinesection that utilize a primary airflow into the engine to generate poweror propel the vehicle. The gas turbine engine is typically mountedwithin a housing, such as a nacelle. A bypass airflow flows through apassage between the housing and the engine and exits from the engine atan outlet.

Presently, conventional thrust reversers are used to generate a reversethrust force to slow forward movement of a vehicle, such as an aircraft.One type of conventional thrust reverser utilizes a moveable door stowednear the rear of the nacelle. After touch-down of the aircraft forlanding, the door moves into the bypass airflow passage to deflect thebypass airflow radially outwards into cascades, or vents, that directthe discharge airflow in a forward direction to slow the aircraft.Although effective, this and other conventional thrust reversers serveonly for thrust reversal and, when in the stowed position fornon-landing conditions, do not provide additional functionality. The useof a variable area fan nozzle (VAFN) has been proposed for low pressureratio fan designs to improve the propulsive efficiency of high bypassratio gas turbine engines. Integrating the VAFN functionality into acommon set of thrust reverser cascades operated by a common actuationsystem represents a significant reduction in complexity and weight.

SUMMARY

A gas turbine engine system according to an exemplary aspect of thepresent disclosure may include a core engine defined about an axis, afan driven by the core engine about the axis to generate bypass flow,and at least one integrated mechanism in communication with the bypassflow. The bypass flow defines a bypass ratio greater than about six (6).The at least one integrated mechanism includes a variable area fannozzle (VAFN) and thrust reverser, and a plurality of positions tocontrol bypass flow.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the bypass flow is arranged to communicatewith an exterior environment when the integrated mechanism is in adeployed position.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the integrated mechanism includes a pluralityof apertures to enable the communication of the bypass flow with theexterior environment when the integrated mechanism is in the deployedposition.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the integrated mechanism includes a singleactuator set to move between the plurality of positions.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the thrust reverser has a stowed position anda deployed position to divert the bypass flow in a thrust reversingdirection.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, a gear system is driven by the core engine.The fan is driven by the gear system. The gear system defines a gearreduction ratio of greater than about 2.3.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, a gear system is driven by the core engine.The fan is driven by the gear system. The gear system defines a gearreduction ratio of greater than 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the core engine includes a low pressureturbine which defines a pressure ratio that is greater than about five(5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the core engine includes a low pressureturbine which defines a pressure ratio that is greater than five (5).

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the at least one integrated mechanism isarranged to change a pressure ratio across the fan.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the bypass ratio is greater than about 10.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, the bypass ratio is greater than 10.

In a further non-limiting embodiment of any of the foregoing gas turbineengine system embodiments, a gear system is driven by the core engine.The fan is driven by the gear system with a gear reduction ratio greaterthan 2.5. The gear system is an epicycle gear train. The core engineincludes a low pressure turbine which defines a pressure ratio that isgreater than five (5).

A gas turbine engine according to another exemplary aspect of thepresent disclosure may include a core engine defined about an axis, afan couple to be driven by said core engine about the axis to generate abypass flow, and at least one integrated mechanism in communication withthe bypass flow. The core engine includes at least a low pressureturbine which defines a pressure ratio that is greater than about five(5). The at least one integrated mechanism includes a variable area fannozzle (VAFN) and a thrust reverser. The integrated mechanism also mayincludes a plurality of positions to control bypass flow. The integratedmechanism includes a section common to the thrust reverser and VAFN.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the integrated mechanism includes at least oneactuator set to move between the plurality of positions.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the thrust reverser has a stowed position and adeployed position to divert the bypass flow in a thrust reversingdirection.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the common section is moveable between a pluralityof axial positions and has a plurality of apertures providing a flowpath for the bypass flow to reach an exterior environment of the gasturbine engine.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, a gear system is included. The core engine drivesthe fan via the gear system, which defines a gear reduction ratio ofgreater than about 2.3.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, a gear system is included. The core engine drivesthe fan via the gear system, which defines a gear reduction ratio ofgreater than 2.5.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow defines a bypass ratio greater thanabout ten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the bypass flow defines a bypass ratio greater thanten (10).

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the thrust reverser includes a blocker door moveablebetween a stowed position and a deployed position and a link having oneend connected to the blocker door and an opposite end connected to asupport.

In a further non-limiting embodiment of any of the foregoing gas turbineengine embodiments, the blocker door includes a slot having a T-shapedcross section, the slot slidably receiving the link.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows.

FIG. 1 illustrates selected portions of an example gas turbine enginesystem having a mechanism that integrates a variable fan nozzleintegrated and a thrust reverser.

FIG. 2 illustrates a perspective view of the example gas turbine enginesystem with cascades exposed for thrust reversal.

FIG. 3A illustrates a schematic view of the mechanism having an axiallymoveable section that is in a closed position.

FIG. 3B illustrates a schematic view of the axially moveable section inan intermediate position for altering a discharge flow from the gasturbine engine.

FIG. 3C illustrates a schematic view of the axially moveable section inan open position for generating a thrust reversing force.

FIG. 4 illustrates a blocker door of the thrust reverser.

FIG. 5 illustrates a view of an example slot of the blocker dooraccording to the section shown in FIG. 4.

DETAILED DESCRIPTION

FIG. 1 illustrates a schematic view of selected portions of an examplegas turbine engine 10 suspended from an engine pylon 12 of an aircraft,as is typical of an aircraft designed for subsonic operation. The gasturbine engine 10 is circumferentially disposed about an enginecenterline, or axial centerline axis A. The gas turbine engine 10includes a fan 14, a low pressure compressor 16 a, a high pressurecompressor 16 b, a combustion section 18, a low pressure turbine 20 a,and a high pressure turbine 20 b. As is well known in the art, aircompressed in the compressors 16 a, 16 b is mixed with fuel that isburned in the combustion section 18 and expanded in the turbines 20 aand 20 b. The turbines 20 a and 20 b are coupled for rotation with,respectively, rotors 22 and 24 (e.g., spools) to rotationally drive thecompressors 16 a, 16 b and the fan 14 in response to the expansion. Inthis example, the rotor 22 also drives the fan 14 through a gear train24.

The engine 10 is preferably a high-bypass geared architecture aircraftengine. In one disclosed, non-limiting embodiment, the engine 10 bypassratio is greater than about six (6) to ten (10), the gear train 22 is anepicyclic gear train such as a planetary gear system or other gearsystem with a gear reduction ratio of greater than about 2.3 and the lowpressure turbine 18 has a pressure ratio that is greater than about 5.In the example shown, the gas turbine engine 10 is a high bypassturbofan arrangement. In one example, the bypass ratio is greater than10, and the fan 14 diameter is substantially larger than the diameter ofthe low pressure compressor 16 a. The low pressure turbine 20 a has apressure ratio that is greater than 5, in one example. The gear train 24is an epicycle gear train, for example, a star gear train, providing agear reduction ratio of greater than 2.5. It should be understood,however, that the above parameters are only exemplary of a contemplatedgeared turbofan engine. That is, the invention is applicable to otherengines.

An outer housing, nacelle 28, (also commonly referred to as a fannacelle) extends circumferentially about the fan 14. A fan bypasspassage 32 extends between the nacelle 28 and an inner housing, innercowl 34, which generally surrounds the compressors 16 a, 16 b andturbines 20 a, 20 b. In this example, the gas turbine engine 10 includesintegrated mechanisms 30 that are coupled to the nacelle 28. Theintegrated mechanisms 30 integrate functions of a variable fan nozzleand a thrust reverser, as will be described below. Any number ofintegrated mechanisms 30 may be used to meet the particular needs of anengine. In this example, two integrated mechanisms 30 are used, one oneach semi-circular half of the nacelle 28.

In operation, the fan 14 draws air into the gas turbine engine 10 as acore flow, C, and into the bypass passage 32 as a bypass air flow, D.The bypass air flow D is discharged as a discharge flow through a rearexhaust 36 associated with the integrated mechanism 30 near the rear ofthe nacelle 28 in this example. The core flow C is discharged from apassage between the inner cowl 34 and a tail cone 38.

For the gas turbine engine 10 shown FIG. 1, a significant amount ofthrust may be provided by the discharge flow due to the high bypassratio. Thrust is a function of density, velocity, and area. One or moreof these parameters can be manipulated to vary the amount and directionof thrust provided or to enhance conditions for aircraft control,operation of the fan 14, operation of other components associated withthe bypass passage 32, or operation of the gas turbine engine 10. Forexample, an effective reduction in area of the rear exhaust 36 causes anair pressure increase within the bypass passage 32 that in turn changesa pressure ratio across the fan 14.

In the disclosed example, the integrated mechanism 30 includes astructure associated with the rear exhaust 36 to change one or more ofthese parameters. However, it should be understood that the bypass flowor discharge flow may be effectively altered by other than structuralchanges, for example, by altering a flow boundary layer. Furthermore, itshould be understood that effectively altering a cross-sectional area ofthe rear exhaust 36 is not limited to physical locations approximate tothe exit of the nacelle 28, but rather, includes altering the bypassflow D by any suitable means.

Referring to FIGS. 2 and 3A, the integrated mechanism 30 in this exampleincludes a nozzle 40 and a thrust reverser 42. The nozzle 40 and thrustreverser include a common part, section 44, which is moveable between aplurality of axial positions relative to the centerline axis A. In thisexample, the section 44 is a hollow sleeve-like structure that extendsabout a cascade section 46. Actuators 48 are mounted within the nacelle28 in this example. Links 50 extend through the cascade section 46 andare coupled on one end with the respective actuators 48 and on anopposite end with the section 44 in a known manner A controller 49communicates with the actuators 48 to selectively axially move thesection 44. The controller 49 may be dedicated to controlling theintegrated mechanism 30, integrated into an existing engine controllerwithin the gas turbine engine 10, or be incorporated with other knownaircraft or engine controls. Alternatively, one or more of the actuators48 are mounted within the cascade section 46 in a known manner

In the disclosed example, the cascade section 46 includes a plurality ofapertures 52, or vents, that provide a flow path between the bypasspassage 32 and the exterior environment of the gas turbine engine 10.The apertures 52 may be formed in any known suitable shape, such as withairfoil shaped vanes 53 between the apertures 52. In this example, theapertures 52 are arranged in circumferential rows about the cascadesection 46. A first set of apertures 52 a near the forward end of thecascade section 46 are angled aft and a second set of apertures 52 b aftof the first set of apertures 52 a are angled forward. Axial movement ofthe section 44 selectively opens, or exposes, the apertures 52 a,apertures 52 b, or both to provide an auxiliary passage for thedischarge flow, as will be described below.

In the illustrated example, there are two circumferential rows in thefirst set of apertures 52 a and a larger number of circumferential rowsin the second set of apertures 52 b. In one example, two circumferentialrows in the first set of apertures 52 a is adequate for altering thedischarge flow, as will be described. However, it is to be understoodthat one circumferential row or greater than two circumferential rowsmay be used for smaller or larger alterations, respectively.

The thrust reverser 42 includes a blocker door 62 having a stowedposition (FIG. 3A) and a fully deployed position (FIG. 3C). The blockerdoor 62 is pivotally connected to the section 44 at connection 63. Adrag link 64 includes one end that is slidably connected to the blockerdoor 62 and an opposite end that is connected to a support, the innercowl 34 in this example. Although only one drag link 64 is shown, it isto be understood that any suitable number of drag links 64 may be used.

Referring to FIGS. 4 and 5, the blocker door 62 includes a slot 66 forslidably connecting the drag link 64 to the blocker door 62. In thisexample, the shape of the slot 66 is adapted to receive and retain theend of the drag link 64. For example, the slot 66 is T-shaped and theend of drag link 64 includes laterally extending slide members 68, suchas rollers, bearings, friction material, or other known suitablemechanism for allowing the end of the drag link 64 to slide along theslot 66. Given this description, one of ordinary skill in the art willrecognize alternative suitable slot shapes or sliding connections tomeet their particular needs.

In operation, the controller 49 selectively commands the actuators 48 tomove the section 44 between the plurality of axial positions to alterthe discharge flow or provide thrust reversal. FIG. 3A illustrates thesection 44 in a first axial position (i.e., a closed position) sealedagainst the nacelle 28. In the closed position, the section 44completely covers the cascade section 46 such that the discharge flowexits axially through the rear exhaust 36.

FIG. 3B illustrates the section 44 in a second axial position spacedapart from the nacelle 28 to provide an opening there between and exposea portion of the cascade section 46. In the second position, the firstset of apertures 52 a are exposed to provide an auxiliary passage forthe discharge flow. The auxiliary passage provides an additional passage(i.e., additional effective cross-sectional flow area) for exit of thedischarge flow from the bypass passage 32 to thereby alter the dischargeflow. A portion of the discharge flow flows through the first set ofapertures 52 a and is directed in the aft direction. Although the aftangle in the illustrated example is not parallel to the centerline axisA, a geometric component of the aft angle is parallel. The geometriccomponent of the discharge flow that is parallel to the centerline axisA provides the benefit of maintaining a portion of the thrust generatedby the discharge flow.

Upon movement of the section 44 between the first position and thesecond position, the blocker door 62 remains in the stowed position. Theconnection between the drag link 64 and the slot 66 provides a range oflost motion movement. That is, the movement of the section 44 causes thedrag link 64 to slide along the slot 66 of the blocker door 62 withoutmoving the blocker door 62 into the deployed position.

FIG. 3C illustrates the section 44 in a third axial position (i.e., athrust reverse position). Movement of the section 44 beyond the secondposition toward the third position causes the end of the drag link 64 toengage an end 70 of the slot 66. Once engaged, the drag link 64 pivotsthe blocker door 62 about the connection 63 and into the bypass passage32. The blocker door 62 deflects the discharge flow radially outwardsrelative to the centerline axis A toward the cascade section 46. Themovement of the section 44 to the third position also exposes theapertures 52 b. The deflected discharge flow enters the second set ofapertures 52 b, which angle the discharge flow in the forward directionto generate a reverse thrust force.

In this example, there are more apertures 52 within the first set ofapertures 52 b than in the second set of apertures 52 a. Thus, thereverse thrust force due to discharge flow through the second set ofapertures 52 b overcomes any thrust due to aft discharge flow from theapertures 52 a.

The disclosed example integrated mechanism 30 thereby integrates thefunction of altering the discharge flow with the thrust reversingfunction. The integrated mechanism 30 utilizes a single set or system ofactuators 48 to eliminate the need for separate actuators or sets ofactuators for altering the discharge flow and deploying the thrustreverser. Using a single actuator or set of actuators 48 as in thedisclosed examples eliminates at least some of the actuators that wouldotherwise be used, thereby reducing the weight of the gas turbine engine10 and increasing the fuel efficiency.

Although a preferred embodiment of this invention has been disclosed, aworker of ordinary skill in this art would recognize that certainmodifications would come within the scope of this invention. For thatreason, the following claims should be studied to determine the truescope and content of this invention.

What is claimed is:
 1. A gas turbine engine comprising: a core enginedefined about an axis; a fan driven by said core engine about said axisto generate bypass flow, wherein the bypass flow defines a bypass ratiogreater than about six (6); and at least one integrated mechanism incommunication with the bypass flow, the at least one integratedmechanism including a variable area fan nozzle (VAFN) and a thrustreverser, the integrated mechanism including a plurality of positions tocontrol bypass flow.
 2. The gas turbine engine of claim 1, wherein thebypass flow is arranged to communicate with an exterior environment whenthe integrated mechanism is in a deployed position.
 3. The gas turbineengine of claim 2, wherein the integrated mechanism includes a pluralityof apertures to enable the communication of the bypass flow with theexterior environment when the integrated mechanism is in the deployedposition.
 4. The gas turbine engine of claim 1, wherein the integratedmechanism includes a single actuator set to move between the pluralityof positions.
 5. The gas turbine engine of claim 1, wherein the thrustreverser has a stowed position and a deployed position to divert thebypass flow in a thrust reversing direction.
 6. The gas turbine engineof claim 1, further comprising a gear system driven by said core engine,said fan driven by said gear system, wherein the gear system defines agear reduction ratio of greater than about 2.3.
 7. The gas turbineengine of claim 1, further comprising a gear system driven by said coreengine, said fan driven by said gear system, wherein the gear systemdefines a gear reduction ratio of greater than 2.5.
 8. The gas turbineengine of claim 1, wherein the core engine includes a low pressureturbine which defines a pressure ratio that is greater than about five(5).
 9. The gas turbine engine of claim 1, wherein the core engineincludes a low pressure turbine which defines a pressure ratio that isgreater than five (5).
 10. The gas turbine engine of claim 1, whereinthe at least one integrated mechanism is arranged to change a pressureratio across the fan.
 11. The gas turbine engine of claim 1, wherein thebypass ratio is greater than about
 10. 12. The gas turbine engine ofclaim 1, wherein the bypass ratio is greater than
 10. 13. The gasturbine engine of claim 1, further comprising a gear system driven bysaid core engine, wherein said fan is driven by the gear system with agear reduction ratio greater than 2.5, wherein said gear system is anepicycle gear train, and wherein the core engine includes a low pressureturbine which defines a pressure ratio that is greater than five (5).14. A gas turbine engine comprising: a core engine defined about anaxis, said core engine including at least a low pressure turbine whichdefines a pressure ratio that is greater than about five (5); a fancouple to be driven by said core engine about said axis to generate abypass flow; and at least one integrated mechanism in communication withthe bypass flow, the at least one integrated mechanism including avariable area fan nozzle (VAFN) and a thrust reverser, the integratedmechanism including a plurality of positions to control bypass flow,wherein the integrated mechanism includes a section common to the thrustreverser and VAFN.
 15. The gas turbine engine of claim 14, wherein theintegrated mechanism includes at least one actuator set to move betweenthe plurality of positions.
 16. The gas turbine engine of claim 14,wherein the thrust reverser has a stowed position and a deployedposition to divert the bypass flow in a thrust reversing direction. 17.The gas turbine engine of claim 14, wherein the common section ismoveable between a plurality of axial positions and has a plurality ofapertures providing a flow path for the bypass flow to reach an exteriorenvironment of the gas turbine engine.
 18. The gas turbine engine ofclaim 14, further including a gear system, wherein the core enginedrives the fan via the gear system, the gear system defines a gearreduction ratio of greater than about 2.3.
 19. The gas turbine engine ofclaim 14, further including a gear system, wherein the core enginedrives the fan via the gear system, the gear system defines a gearreduction ratio of greater than 2.5.
 20. The gas turbine engine of claim14, wherein the bypass flow defines a bypass ratio greater than aboutten (10).
 21. The gas turbine engine of claim 14, wherein the bypassflow defines a bypass ratio greater than ten (10).
 22. The gas turbineengine of claim 14, wherein the thrust reverser includes a blocker doormoveable between a stowed position and a deployed position and a linkhaving one end connected to the blocker door and an opposite endconnected to a support.
 23. The gas turbine engine of claim 22, whereinthe blocker door includes a slot having a T-shaped cross section, theslot slidably receiving the link.